US5112009A - Pitch control system for aircraft - Google Patents

Pitch control system for aircraft Download PDF

Info

Publication number
US5112009A
US5112009A US07/690,310 US69031091A US5112009A US 5112009 A US5112009 A US 5112009A US 69031091 A US69031091 A US 69031091A US 5112009 A US5112009 A US 5112009A
Authority
US
United States
Prior art keywords
incidence
nose
aerodynamic
aircraft
horizontal plane
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US07/690,310
Inventor
Jacques Farineau
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Airbus Operations SAS
Airbus Group SAS
Original Assignee
Airbus Group SAS
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Airbus Group SAS filed Critical Airbus Group SAS
Assigned to AEROSPATIALE SOCIETE NATIONALE INDUSTRIELLE reassignment AEROSPATIALE SOCIETE NATIONALE INDUSTRIELLE ASSIGNMENT OF ASSIGNORS INTEREST. Assignors: FARINEAU, JACQUES
Application granted granted Critical
Publication of US5112009A publication Critical patent/US5112009A/en
Assigned to AIRBUS FRANCE reassignment AIRBUS FRANCE ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: EUROPEAN AERONAUTIC DEFENCE AND SPACE COMPANY EADS FRANCE
Assigned to EUROPEAN AERONAUTIC DEFENCE AND SPACE COMPANY EADS FRANCE reassignment EUROPEAN AERONAUTIC DEFENCE AND SPACE COMPANY EADS FRANCE CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: AEROSPATIALE SOCIETE NATIONALE INDUSTRIELLE
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/24—Transmitting means
    • B64C13/38—Transmitting means with power amplification
    • B64C13/50—Transmitting means with power amplification using electrical energy
    • B64C13/503—Fly-by-Wire
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/02—Initiating means
    • B64C13/16—Initiating means actuated automatically, e.g. responsive to gust detectors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/24—Transmitting means
    • B64C13/38—Transmitting means with power amplification
    • B64C13/50—Transmitting means with power amplification using electrical energy
    • B64C13/506—Transmitting means with power amplification using electrical energy overriding of personal controls; with automatic return to inoperative position
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability

Definitions

  • the present invention relates to a pitch control system for aircraft flying at a Mach number greater than 0.7.
  • the pitching aerodynamic couple on the aircraft varies as a function of its aerodynamic incidence in such a manner that:
  • said aerodynamic pitching couple increases with increasing aerodynamic incidence and decreases with decreasing aerodynamic incidence
  • said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence.
  • An object of the present invention is to remedy this drawback for an aircraft having a horizontal plane of adjustable deflection and air brakes that tend to pitch the aircraft nose down when in the deployed position.
  • the present invention provides a pitch control system for an aircraft flying at a Mach number greater than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes that have a nose-down effect, the aerodynamic pitching couple on said aircraft varying as a function of its aerodynamic incidence in such a manner that:
  • said aerodynamic pitching couple increases with increasing aerodynamic incidence and decreases with decreasing aerodynamic incidence
  • said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence
  • system comprises:
  • first means providing the aerodynamic incidence of said aircraft at each instant
  • third means for generating a signal representative of the deflection response rate of said adjustable horizontal plane in response to said first command
  • the tendency of the aircraft to pitch suddenly nose up on exceeding said first incidence threshold is compensated by a nose-down command applied to the adjustable horizontal plane and to the air brakes that produce a nose-down effect. If the aircraft pitches nose up too quickly for it to be possible to provide compensation instantly by means of said adjustable horizontal plane (which operates relatively slowly), said air brakes instantly provide additional compensation since they can be deployed quickly. It is thus preferable to use the adjustable horizontal plane to eliminate inversions in the way aerodynamic pitching couple varies as a function of aerodynamic incidence, but if said adjustable horizontal plane is not fast enough, then the air brakes provide the required effect temporarily.
  • phase advance component may be proportional to the pitching rate of the aircraft. In a variant it may be proportional to the derivative of the aerodynamic incidence of the aircraft.
  • said second means are constituted by a table causing values of the aerodynamic incidence of the aircraft to correspond to deflection values for said adjustable horizontal plane, said deflection values being zero when the aerodynamic incidence is less than said first threshold.
  • a device is included between said second means and said adjustable horizontal plane to limit the rate at which said first command varies as a function of the real deflection rate capability of said adjustable horizontal plane.
  • said third means are constituted by a simulator device for simulating the deflection response rate of said adjustable horizontal plane, said third means receiving said first command.
  • a simulator device may be constituted, for example, by calculation or tabulation means incorporating the transfer function of said adjustable horizontal plane.
  • said fifth means are constituted by a table causing positive values of said difference to correspond to deflection values of said air brakes having a nose-down effect, said deflection values being zero for zero or negative values of said difference.
  • said aircraft when said aircraft also includes air brakes that have a nose-up pitching effect, it is advantageous for the system of the invention to include sixth means responsive to said difference to generate a nose-up third command for said air brakes having a nose-up effect.
  • said sixth means are constituted by a table causing negative values of said difference to correspond to deflection values for said air brakes having a nose-up effect, said deflection values being zero for zero or positive values of said difference.
  • the present invention also provides an aircraft for flying at a Mach number higher than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes having a nose-down effect, said aircraft including a pitch control system as specified above.
  • FIG. 1 is a diagrammatic perspective view of an aircraft suitable for being improved by the present invention
  • FIG. 2 is a graph showing how the aerodynamic pitching couple on said aircraft varies as a function of its aerodynamic incidence
  • FIG. 3 is a graph similar to FIG. 2 showing how the system of the present invention operates.
  • FIG. 4 is a block diagram of an embodiment of the present invention.
  • the large-capacity civil aircraft 1 shown in FIG. 1 is capable of flying at high Mach numbers (greater than 0.7) and it includes numerous moving aerodynamic control surfaces (flaps, ailerons, etc. . . . ) on its wings 2 and 3, including air brakes 4 and air brakes 5.
  • air brakes 4 when the air brakes 4 are in their deployed position they also have a nose-up effect, whereas when the air brakes 5 are in their deployed position they have a nose-down effect in addition to their braking effect.
  • the tail plane 6 on which the elevator 7 is mounted is controllable in deflection, i.e. it is made in the form of an adjustable horizontal plane or stabilizer.
  • the deployment speed of the nose-up effect air brakes 4 and the nose-down effect air brakes 5 is very high (it is not less than 30°/sec), however these air brakes 4 and 5 cannot be used for extended periods for their nose-up or nose-down properties (except during braking), since aircraft performance is greatly degraded by their braking action, and in addition, actuating them gives rise to vibrations which constitute a source of discomfort for passengers.
  • the speed at which the adjustable horizontal plane 6 can be deflected is low, e.g. about 0.5°/sec, such that the adjustable horizontal plane 6 cannot be immediately effective in compensating a sudden change in the incidence of the aircraft 1.
  • zone I where the aerodynamic incidence ⁇ is less than the first threshold ⁇ 0, variation in the aerodynamic pitching couple Cm is at least more or less inversely proportional to the aerodynamic incidence ⁇ .
  • the couple Cm decreases, i.e. the aircraft nose goes down, thereby causing the incidence to decrease.
  • the incidence ⁇ decreases, then the couple Cm increases, the aircraft nose goes up and the incidence increases.
  • flight of the aircraft 1 is thus stable with any departure from equilibrium creating an effect that tends to return the aircraft to equilibrium.
  • the zone I thus corresponds to the domain of normal flight for the aircraft 1 in which the aircraft is longitudinally stable.
  • zone III where the aerodynamic incidence ⁇ is greater than the second threshold ⁇ 2, variation in the aerodynamic pitch couple Cm is again substantially inversely proportional to the aerodynamic incidence ⁇ , as in the zone I, but with greater sensitivity since a small change in incidence corresponds to a larger change in the couple Cm.
  • the zone III thus also corresponds to a domain of stable flight.
  • the curve K1 has a bulge B disposed astride said second threshold ⁇ 2 and representing a double inversion of the way in which the couple Cm varies as a function of the incidence ⁇ .
  • the aircraft 1 encounters atmospheric turbulence that has the effect either of increasing the incidence, or the Mach number, or both, such that the incidence ⁇ becomes greater than the first threshold ⁇ 0.
  • the aircraft 1 then tends to start pitching as in a) above.
  • the load factor applied to the aircraft 1 during such pitching takes on high values that may exceed 2 g, thereby giving rise to considerable passenger discomfort.
  • An object of the present invention is to avoid this discomfort by improving the behavior of the aircraft 1 on going through the bulge B.
  • the invention makes it possible to extend the zone I portion K.I of the curve K past the bulge B.
  • the extension P obtained in this way (dashed line) enables said portion K.I of the curve to meet the portion of the curve K that lies in the zone III, thereby short circuiting the bulge B.
  • a quantity ⁇ Cm is subtracted from the nose-up couple Cm at each value of the aerodynamic incidence ⁇ that is greater than the first threshold ⁇ 0, in such a manner that (Cm- ⁇ Cm) varies along the extension P.
  • the bulge B is thus eliminated and can be made transparent for the pilot.
  • the aircraft 1 is thus stabilized in incidence, even at values of incidence that are greater than the first threshold ⁇ 0.
  • FIG. 4 is a block diagram of an implementation of a system of the present invention suitable for obtaining the operation shown in FIG. 3.
  • This system includes a probe 10 for delivering the aerodynamic incidence ⁇ of the aircraft 1.
  • An adder 11 adds a phase advance component a to this aerodynamic incidence ⁇ .
  • the component a may be proportional to the pitch rate (or speed), or else to the derivative of the aerodynamic incidence ⁇ , and it is generated by a device that is not shown.
  • the aerodynamic incidence ⁇ plus the phase advance component a is transmitted to a table look-up device 12 which provides values ⁇ iHc that correspond to the values ( ⁇ +a) as follows:
  • ⁇ iHc is equal to the nose-down deflection that the adjustable horizontal stabilizer 6 needs to take up to compensate for the nose-up effect applied to the aircraft 1 by virtue of the incidence threshold ⁇ 0 being exceeded.
  • the values ⁇ iHc thus constitute commands for the adjustable horizontal stabilizer 6.
  • the values ⁇ iHc are applied to said adjustable horizontal stabilizer 6 via a limiter device 13 which matches the rate of change in the signal ⁇ iHc to the real capabilities of the adjustable horizontal stabilizer 6.
  • the device 13 converts the signal ⁇ iHc into a signal ⁇ iH which varies at a rate that is adapted to the adjustable horizontal stabilizer 6, and which is applied thereto.
  • the system of the invention includes a device 14 which delivers a signal ⁇ iHe representative of the instantaneous nose-down deflection position occupied by the adjustable horizontal stabilizer 6 in response to the signal ⁇ iH.
  • the device 14 may be a simulator device incorporating the transfer function of the adjustable horizontal stabilizer 6 and providing a value of the signal ⁇ iHe corresponding to each value of the signal ⁇ iH.
  • the device 14 may be constituted by a table look-up device including such a correspondence table.
  • a subtractor 15 subtracts the signal ⁇ iHe generated by the device 14 from the signal ⁇ iHc generated by the table look-up device 12. It may be observed that the difference d( ⁇ iH) obtained in this way corresponds to the instantaneous lack of nose-down effect from the adjustable horizontal stabilizer 6 for compensating the nose-up effect on the aircraft 1, which lack is due to the slowness with which said adjustable horizontal stabilizer 6 moves.
  • said difference d( ⁇ iH) is applied to the nose-down effect air brakes 5 via a device 16 which transforms said difference into a nose-down deflection command therefor.
  • the device 16 includes a table which provides a nose-down deflection value for the air brakes 5 corresponding to any positive value of said difference, and which provides zero nose-down deflection values if said difference is zero or negative.
  • the nose-up effect on the aircraft 1 is compensated mainly by the nose-down effect of the deflection applied to the adjustable horizontal stabilizer 6;
  • the nose-down effect air brakes 5 provide a nose-down effect to make up for the instantaneous lack of nose-down effect due to the adjustable horizontal stabilizer 6 moving at a limited speed;
  • the difference d( ⁇ iH) may be negative when the incidence ⁇ returns from values greater than ⁇ 0 to values less than ⁇ 0, with the adjustable horizontal stabilizer 6 being in a nose-down position. If such a return to values of ⁇ less than ⁇ 0 takes place quickly, it is necessary to return the adjustable horizontal stabilizer 6 fairly fast to its initial position (prior to taking up a position having a nose-down effect). However, because of its operating inertia, the adjustable horizontal stabilizer 6 cannot return quickly to its initial position. Under such circumstances, a transient lack of nose-up effect appears in the zone I of the curve K.
  • a device 17 which receives the signal d( ⁇ iH) and which generates a nose-up deflection command for the nose-up effect air brakes 4.
  • the device 17 includes a table causing such negative values of said difference d( ⁇ iH) to correspond to nose-up command values, with such nose-up command values being zero whenever said difference is positive or zero.

Landscapes

  • Engineering & Computer Science (AREA)
  • Automation & Control Theory (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
  • Traffic Control Systems (AREA)
  • Radar Systems Or Details Thereof (AREA)
  • Mechanical Control Devices (AREA)

Abstract

A pitch control system for an aircraft flying at a Mach number greater than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes having a nose-down effect. According to the invention, the system comprises:
first means providing the aerodynamic incidence (α) of said aircraft at each instant;
second means for generating a nose-down first command (ΔiHC) for said adjustable horizontal plane when said aerodynamic incidence is greater than said first threshold (α0), the amplitude of said first command being such that if said adjustable horizontal plane were to take up the corresponding position instantaneously, it would compensate the increase in the nose-up effect on said aircraft due to the way in which said aerodynamic pitching couple varies above said first threshold;
third means for generating a signal (ΔiHe) representative of the deflection response rate of said adjustable horizontal plane in response to said first command;
fourth means for forming the difference d(ΔiH) between said first command (ΔiHc) and said signal (ΔiHe) representative of the response of said adjustable horizontal plane; and
fifth means for responding to said difference d(ΔiH) to generate a nose-down second command for said air brakes having a nose-down effect.

Description

The present invention relates to a pitch control system for aircraft flying at a Mach number greater than 0.7.
BACKGROUND OF THE INVENTION
Under such circumstances, the pitching aerodynamic couple on the aircraft varies as a function of its aerodynamic incidence in such a manner that:
below a first incidence threshold said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence;
between said first incidence threshold and a second incidence threshold greater than said first threshold, said aerodynamic pitching couple increases with increasing aerodynamic incidence and decreases with decreasing aerodynamic incidence; and
above said second incidence threshold said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence.
Consequently, if the aerodynamic incidence of the aircraft accidentally exceeds said first threshold, the aircraft enters a domain of unstable flight in which any increase in incidence causes the aerodynamic pitching couple to increase which in turn increases the angle of incidence. As a result, the aircraft suddenly pitches nose up, and this may be followed by the aircraft suddenly pitching nose down when said second threshold is exceeded, thus producing great discomfort for the passengers of the aircraft.
An object of the present invention is to remedy this drawback for an aircraft having a horizontal plane of adjustable deflection and air brakes that tend to pitch the aircraft nose down when in the deployed position.
SUMMARY OF THE INVENTION
To this end, the present invention provides a pitch control system for an aircraft flying at a Mach number greater than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes that have a nose-down effect, the aerodynamic pitching couple on said aircraft varying as a function of its aerodynamic incidence in such a manner that:
below a first incidence threshold said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence;
between said first incidence threshold and a second incidence threshold greater than said first threshold, said aerodynamic pitching couple increases with increasing aerodynamic incidence and decreases with decreasing aerodynamic incidence; and
above said second incidence threshold said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence;
wherein the system comprises:
first means providing the aerodynamic incidence of said aircraft at each instant;
second means for generating a nose-down first command for said adjustable horizontal plane when said aerodynamic incidence is greater than said first threshold, the amplitude of said first command being such that if said adjustable horizontal plane were to take up the corresponding position instantaneously, it would compensate the increase in the nose-up effect on said aircraft due to the way in which said aerodynamic pitching couple varies above said first threshold;
third means for generating a signal representative of the deflection response rate of said adjustable horizontal plane in response to said first command;
fourth means for forming the difference between said first command and said signal representative of the response of said adjustable horizontal plane; and
fifth means for responding to said difference to generate a nose-down second command for said air brakes having a nose-down effect.
Thus, according to the invention, the tendency of the aircraft to pitch suddenly nose up on exceeding said first incidence threshold is compensated by a nose-down command applied to the adjustable horizontal plane and to the air brakes that produce a nose-down effect. If the aircraft pitches nose up too quickly for it to be possible to provide compensation instantly by means of said adjustable horizontal plane (which operates relatively slowly), said air brakes instantly provide additional compensation since they can be deployed quickly. It is thus preferable to use the adjustable horizontal plane to eliminate inversions in the way aerodynamic pitching couple varies as a function of aerodynamic incidence, but if said adjustable horizontal plane is not fast enough, then the air brakes provide the required effect temporarily.
In order to anticipate the nose-up correction applied to the aircraft, it is preferable to provide an adder between said first and second means, the adder adding a phase advance component to said aerodynamic incidence.
Such a phase advance component may be proportional to the pitching rate of the aircraft. In a variant it may be proportional to the derivative of the aerodynamic incidence of the aircraft.
Advantageously, said second means are constituted by a table causing values of the aerodynamic incidence of the aircraft to correspond to deflection values for said adjustable horizontal plane, said deflection values being zero when the aerodynamic incidence is less than said first threshold.
In order to take account of the speed with which the adjustable horizontal plane can be deployed, a device is included between said second means and said adjustable horizontal plane to limit the rate at which said first command varies as a function of the real deflection rate capability of said adjustable horizontal plane.
Further, said third means are constituted by a simulator device for simulating the deflection response rate of said adjustable horizontal plane, said third means receiving said first command. Such a simulator device may be constituted, for example, by calculation or tabulation means incorporating the transfer function of said adjustable horizontal plane.
Preferably, said fifth means are constituted by a table causing positive values of said difference to correspond to deflection values of said air brakes having a nose-down effect, said deflection values being zero for zero or negative values of said difference.
As described below, when said aircraft also includes air brakes that have a nose-up pitching effect, it is advantageous for the system of the invention to include sixth means responsive to said difference to generate a nose-up third command for said air brakes having a nose-up effect. Advantageously, said sixth means are constituted by a table causing negative values of said difference to correspond to deflection values for said air brakes having a nose-up effect, said deflection values being zero for zero or positive values of said difference.
The present invention also provides an aircraft for flying at a Mach number higher than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes having a nose-down effect, said aircraft including a pitch control system as specified above.
BRIEF DESCRIPTION OF THE DRAWINGS
An embodiment of the invention is described by way of example with reference to the accompanying drawings, in which:
FIG. 1 is a diagrammatic perspective view of an aircraft suitable for being improved by the present invention;
FIG. 2 is a graph showing how the aerodynamic pitching couple on said aircraft varies as a function of its aerodynamic incidence;
FIG. 3 is a graph similar to FIG. 2 showing how the system of the present invention operates; and
FIG. 4 is a block diagram of an embodiment of the present invention.
When identical references are used in more than one of the figures, they designate items that are similar.
MORE DETAILED DESCRIPTION
The large-capacity civil aircraft 1 shown in FIG. 1 is capable of flying at high Mach numbers (greater than 0.7) and it includes numerous moving aerodynamic control surfaces (flaps, ailerons, etc. . . . ) on its wings 2 and 3, including air brakes 4 and air brakes 5. In addition to their braking effect, when the air brakes 4 are in their deployed position they also have a nose-up effect, whereas when the air brakes 5 are in their deployed position they have a nose-down effect in addition to their braking effect. In addition, the tail plane 6 on which the elevator 7 is mounted is controllable in deflection, i.e. it is made in the form of an adjustable horizontal plane or stabilizer.
In conventional manner, the deployment speed of the nose-up effect air brakes 4 and the nose-down effect air brakes 5 is very high (it is not less than 30°/sec), however these air brakes 4 and 5 cannot be used for extended periods for their nose-up or nose-down properties (except during braking), since aircraft performance is greatly degraded by their braking action, and in addition, actuating them gives rise to vibrations which constitute a source of discomfort for passengers.
Also in conventional manner, the speed at which the adjustable horizontal plane 6 can be deflected is low, e.g. about 0.5°/sec, such that the adjustable horizontal plane 6 cannot be immediately effective in compensating a sudden change in the incidence of the aircraft 1.
If, as shown in the graph of FIG. 2, variation in the aerodynamic pitching couple Cm on the entire aircraft 1 at a fixed position of the adjustable horizontal plane 6 is investigated as a function of the aerodynamic incidence α of said aircraft, then a curve K is obtained similar to the curve K1 of FIG. 2, which curve comprises three zones I, II, and III separated by incidence thresholds α0 and α2 and corresponding to different types of behavior for the aircraft 1.
In the zone I, where the aerodynamic incidence α is less than the first threshold α0, variation in the aerodynamic pitching couple Cm is at least more or less inversely proportional to the aerodynamic incidence α. Thus, if the incidence α increases, then the couple Cm decreases, i.e. the aircraft nose goes down, thereby causing the incidence to decrease. Conversely, if the incidence α decreases, then the couple Cm increases, the aircraft nose goes up and the incidence increases. In the zone I, flight of the aircraft 1 is thus stable with any departure from equilibrium creating an effect that tends to return the aircraft to equilibrium. The zone I thus corresponds to the domain of normal flight for the aircraft 1 in which the aircraft is longitudinally stable.
In the zone II, where the aerodynamic incidence α is greater than the first threshold α0 but less than the second threshold α2, variation in the aerodynamic pitching couple Cm becomes at least more or less directly proportional to the aerodynamic incidence α. Consequently, if the incidence α increases, then the couple Cm increases, thus raising the nose of the aircraft which has the effect of further increasing the incidence. Conversely, if the incidence α decreases, then the couple Cm also decreases such that the nose of the aircraft 1 goes down, thereby further decreasing the incidence. Under such circumstances, pitching motion of the aircraft 1 is divergent, with any departure from equilibrium creating an effect which tends to move the aircraft 1 further from said equilibrium.
Finally, in the zone III, where the aerodynamic incidence α is greater than the second threshold α2, variation in the aerodynamic pitch couple Cm is again substantially inversely proportional to the aerodynamic incidence α, as in the zone I, but with greater sensitivity since a small change in incidence corresponds to a larger change in the couple Cm. The zone III thus also corresponds to a domain of stable flight.
Thus, because of the zone of instability II, the curve K1 has a bulge B disposed astride said second threshold α2 and representing a double inversion of the way in which the couple Cm varies as a function of the incidence α.
This double inversion is due to a phenomenon which is purely aerodynamic. It appears at a high Mach number (greater than 0.7) and the higher the Mach number at which it appears the more pronounced the effect. Further, the higher the Mach number, the lower the first incidence threshold α0. This is illustrated in FIG. 2 which has a curve K2 analogous to the curve K1, but corresponding to a Mach number M2 which is lower than the Mach number M1 associated with the curve K1.
In cruising flight of the aircraft 1, its incidences are selected to lie in the zone I below the first threshold α0, such that the double inversion in the way the couple Cm varies generally does not occur during flight of said aircraft. However, when cruising at a high Mach number, circumstances may arise in which this effect disturbs flight. For example, under manual control:
a) the pilot of the aircraft 1 pulls on the control stick, thereby increasing the incidence so that it exceeds the first threshold α0. As a result the aircraft 1 pitches suddenly nose up (zone II) until the incidence reaches the second threshold α2, whereupon it pitches just as suddenly nose down (zone III); or
b) the aircraft 1 encounters atmospheric turbulence that has the effect either of increasing the incidence, or the Mach number, or both, such that the incidence α becomes greater than the first threshold α0. The aircraft 1 then tends to start pitching as in a) above.
In either case, the load factor applied to the aircraft 1 during such pitching takes on high values that may exceed 2 g, thereby giving rise to considerable passenger discomfort.
An object of the present invention is to avoid this discomfort by improving the behavior of the aircraft 1 on going through the bulge B.
To do this, and as illustrated by the graph of FIG. 3, the invention makes it possible to extend the zone I portion K.I of the curve K past the bulge B. The extension P obtained in this way (dashed line) enables said portion K.I of the curve to meet the portion of the curve K that lies in the zone III, thereby short circuiting the bulge B. In other words, a quantity ΔCm is subtracted from the nose-up couple Cm at each value of the aerodynamic incidence α that is greater than the first threshold α0, in such a manner that (Cm-ΔCm) varies along the extension P. The bulge B is thus eliminated and can be made transparent for the pilot. The aircraft 1 is thus stabilized in incidence, even at values of incidence that are greater than the first threshold α0.
FIG. 4 is a block diagram of an implementation of a system of the present invention suitable for obtaining the operation shown in FIG. 3. This system includes a probe 10 for delivering the aerodynamic incidence α of the aircraft 1. An adder 11 adds a phase advance component a to this aerodynamic incidence α. The component a may be proportional to the pitch rate (or speed), or else to the derivative of the aerodynamic incidence α, and it is generated by a device that is not shown.
The aerodynamic incidence α plus the phase advance component a is transmitted to a table look-up device 12 which provides values ΔiHc that correspond to the values (α+a) as follows:
if (α+a) is less than α0, then ΔiHc is zero;
if (α+a) is greater than α0, then ΔiHc is equal to the nose-down deflection that the adjustable horizontal stabilizer 6 needs to take up to compensate for the nose-up effect applied to the aircraft 1 by virtue of the incidence threshold α0 being exceeded.
The values ΔiHc thus constitute commands for the adjustable horizontal stabilizer 6. In order to take account of the slowness with which the stabilizer moves, the values ΔiHc are applied to said adjustable horizontal stabilizer 6 via a limiter device 13 which matches the rate of change in the signal ΔiHc to the real capabilities of the adjustable horizontal stabilizer 6. Thus, the device 13 converts the signal ΔiHc into a signal δiH which varies at a rate that is adapted to the adjustable horizontal stabilizer 6, and which is applied thereto.
In addition, the system of the invention includes a device 14 which delivers a signal ΔiHe representative of the instantaneous nose-down deflection position occupied by the adjustable horizontal stabilizer 6 in response to the signal δiH. The device 14 may be a simulator device incorporating the transfer function of the adjustable horizontal stabilizer 6 and providing a value of the signal ΔiHe corresponding to each value of the signal δiH. The device 14 may be constituted by a table look-up device including such a correspondence table.
A subtractor 15 subtracts the signal ΔiHe generated by the device 14 from the signal ΔiHc generated by the table look-up device 12. It may be observed that the difference d(ΔiH) obtained in this way corresponds to the instantaneous lack of nose-down effect from the adjustable horizontal stabilizer 6 for compensating the nose-up effect on the aircraft 1, which lack is due to the slowness with which said adjustable horizontal stabilizer 6 moves.
Thus, in order to compensate for the slowness in the reaction time of the adjustable horizontal stabilizer 6, said difference d(ΔiH) is applied to the nose-down effect air brakes 5 via a device 16 which transforms said difference into a nose-down deflection command therefor. The device 16 includes a table which provides a nose-down deflection value for the air brakes 5 corresponding to any positive value of said difference, and which provides zero nose-down deflection values if said difference is zero or negative.
With the system of the invention, it can be seen that:
the nose-up effect on the aircraft 1 is compensated mainly by the nose-down effect of the deflection applied to the adjustable horizontal stabilizer 6;
the nose-down effect air brakes 5 provide a nose-down effect to make up for the instantaneous lack of nose-down effect due to the adjustable horizontal stabilizer 6 moving at a limited speed; and
the action of the nose-down effect air brakes 5 falls off progressively as the adjustable horizontal stabilizer 6 moves into position.
It may be observed that the difference d(ΔiH) may be negative when the incidence α returns from values greater than α0 to values less than α0, with the adjustable horizontal stabilizer 6 being in a nose-down position. If such a return to values of α less than α0 takes place quickly, it is necessary to return the adjustable horizontal stabilizer 6 fairly fast to its initial position (prior to taking up a position having a nose-down effect). However, because of its operating inertia, the adjustable horizontal stabilizer 6 cannot return quickly to its initial position. Under such circumstances, a transient lack of nose-up effect appears in the zone I of the curve K. To eliminate this effect, a device 17 is provided which receives the signal d(ΔiH) and which generates a nose-up deflection command for the nose-up effect air brakes 4. The device 17 includes a table causing such negative values of said difference d(ΔiH) to correspond to nose-up command values, with such nose-up command values being zero whenever said difference is positive or zero.

Claims (9)

I claim:
1. A pitch control system for an aircraft flying at a Mach number greater than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes that have a nose-down effect, the aerodynamic pitching couple (Cm) on said aircraft varying as a function of its aerodynamic incidence (α) in such a manner that:
below a first incidence threshold (α0) said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence;
between said first incidence threshold (α0) and a second incidence threshold (α2) greater than said first threshold, said aerodynamic pitching couple increases with increasing aerodynamic incidence and decreases with decreasing aerodynamic incidence; and
above said second incidence threshold (α2) said aerodynamic pitching couple decreases with increasing aerodynamic incidence and increases with decreasing aerodynamic incidence;
wherein the system comprises:
first means providing the aerodynamic incidence (α) of said aircraft at each instant;
an adder for adding a phase advance component (a) to said aerodynamic incidence (α);
second means for generating a nose-down first command (ΔiHc) for said adjustable horizontal plane when said aerodynamic incidence is greater than said first threshold (α0), the amplitude of said first command being such that if said adjustable horizontal plane were to take up the corresponding position instantaneously, it would compensate the increase in the nose-up effect on said aircraft due to the way in which said aerodynamic pitching couple varies above said first threshold;
third means for generating a signal (ΔiHe) representative of the deflection response rate of said adjustable horizontal plane in response to said first command;
fourth means for forming the difference d(ΔiH) between said first command (ΔiHc) and said signal (ΔiHe) representative of the response of said adjustable horizontal plane; and
fifth means for responding to said difference d(ΔiH) to generate a nose-down second command for said air brakes having a nose-down effect.
2. A system according to claim 1, wherein said fifth means are constituted by a table causing positive values of said difference d(ΔiH) to correspond to deflection values of said air brakes having a nose-down effect, said deflection values being zero for zero or negative values of said difference d(ΔiH).
3. A system according to claim 1, wherein said phase advance component is proportional to the pitching rate of the aircraft.
4. A system according to claim 1, wherein said phase advance component is proportional to the derivative of the aerodynamic incidence of the aircraft.
5. A system according to claim 1, wherein said second means are constituted by a table causing values of the aerodynamic incidence (α) of the aircraft to correspond to deflection values for said adjustable horizontal plane, said deflection values being zero when the aerodynamic incidence (α) is less than said first threshold (α0).
6. A system according to claim 1, including a device between said second means and said adjustable horizontal plane to limit the rate at which said first command varies as a function of the real deflection rate capability of said adjustable horizontal plane.
7. A system according to claim 1, wherein said third means are constituted by a device for simulating the deflection response rate of said adjustable horizontal plane, said third means receiving said first command.
8. A system according to claim 1, for use in an aircraft further including air brakes having a nose-up effect, the system including sixth means responsive to said difference d(ΔiH) to generate a nose-up third command for said air brakes having a nose-up effect.
9. A system according to claim 8, wherein said sixth means are constituted by a table causing negative values of said difference d(ΔiH) to correspond to deflection values for said air brakes having a nose-up effect, said deflection values being zero for zero or positive values of said difference d(ΔiH).
US07/690,310 1990-04-24 1991-04-24 Pitch control system for aircraft Expired - Lifetime US5112009A (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR9005227A FR2661149B1 (en) 1990-04-24 1990-04-24 SYSTEM FOR THE PILOTAGE OF AN AIRPLANE IN TANGAGE.
FR9005227 1990-04-24

Publications (1)

Publication Number Publication Date
US5112009A true US5112009A (en) 1992-05-12

Family

ID=9396042

Family Applications (1)

Application Number Title Priority Date Filing Date
US07/690,310 Expired - Lifetime US5112009A (en) 1990-04-24 1991-04-24 Pitch control system for aircraft

Country Status (6)

Country Link
US (1) US5112009A (en)
EP (1) EP0454549B1 (en)
CA (1) CA2040874C (en)
DE (1) DE69101154T2 (en)
ES (1) ES2051089T3 (en)
FR (1) FR2661149B1 (en)

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5375793A (en) * 1992-08-14 1994-12-27 Aerospatiale Societe Nationale Industrielle Process for the control of the control surfaces of an aircraft for the low speed compensation of a lateral path deviation
US5779191A (en) * 1996-11-12 1998-07-14 Brislawn; Mark G. Pylon flap for increasing negative pitching moments
US6641086B2 (en) * 2001-08-14 2003-11-04 Northrop Grumman Corporation System and method for controlling an aircraft
US20050242234A1 (en) * 2004-04-29 2005-11-03 The Boeing Company Lifters, methods of flight control and maneuver load alleviation
US20050263644A1 (en) * 2004-05-28 2005-12-01 Airbus France Control system for airbrakes of an aircraft
WO2006082305A2 (en) * 2005-02-04 2006-08-10 Airbus France Method and device for piloting a pitching aircraft
US20070057114A1 (en) * 2004-07-16 2007-03-15 Airbus France Procedure and device for improving the maneuverability of an aircraft during the approach to landing and flare-out phases
US20080001028A1 (en) * 2000-02-14 2008-01-03 Greg Kendall Aircraft control system
US20080272242A1 (en) * 2005-11-29 2008-11-06 Airbus France Method for ensuring the safety of an aircraft flying horizontally at low speed
WO2008070944A3 (en) * 2006-12-11 2009-04-02 Embraer Aeronautica Sa Flight control system
US20100042271A1 (en) * 2007-03-15 2010-02-18 Torsten Holzhausen Method and device for moveable tail trimming in an aircraft
RU2445671C2 (en) * 2010-02-25 2012-03-20 Московский государственный университет приборостроения и информатики System for adaptive control of aeroplane on pitch angle
US20140236399A1 (en) * 2013-02-19 2014-08-21 Airbus Operations (Sas) Method and device for estimating an unwanted pitch moment of an aircraft, and applications to the pitch control of the aircraft
CN105584627A (en) * 2014-11-12 2016-05-18 波音公司 Method and apparatus to control aircraft horizontal stabilizer
US20180101181A1 (en) * 2016-10-11 2018-04-12 Airbus Operations Sas Aircraft flight control method and system
EP2506107B1 (en) * 2011-03-31 2020-03-11 Honeywell International Inc. System for controlling the speed of an aircraft
CN114056551A (en) * 2022-01-12 2022-02-18 中国空气动力研究与发展中心低速空气动力研究所 Virtual wing belly flap and wing body fusion airplane, constant air blowing method and variable-angle air blowing method

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2719548B1 (en) * 1994-05-03 1996-07-05 Aerospatiale Transport plane with front tail.
CN108628342A (en) * 2018-04-28 2018-10-09 长光卫星技术有限公司 The automatic flight control system and method for unmanned vehicle

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3142457A (en) * 1962-07-16 1964-07-28 Boeing Co Stall pattern lift regulator for airplanes
US3167276A (en) * 1961-09-29 1965-01-26 Honeywell Inc Control apparatus
US3942746A (en) * 1971-12-27 1976-03-09 General Dynamics Corporation Aircraft having improved performance with beaver-tail afterbody configuration
US4003533A (en) * 1973-10-01 1977-01-18 General Dynamics Corporation Combination airbrake and pitch control device
US4142699A (en) * 1977-02-04 1979-03-06 Boeing Commercial Airplane Company Lateral control system
US4261537A (en) * 1979-02-28 1981-04-14 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Velocity vector control system augmented with direct lift control
US4442490A (en) * 1980-09-26 1984-04-10 S-Tec Corporation Aircraft pitch stabilization apparatus
US4485446A (en) * 1981-09-08 1984-11-27 The Boeing Company Aircraft lift control system with acceleration and attitude limiting
US4569494A (en) * 1982-12-23 1986-02-11 The Boeing Company Pitch control of swept wing aircraft
US4744532A (en) * 1984-01-09 1988-05-17 Societe Nationale Industrielle Et Aerospatiale Aircraft flight control system
US4956780A (en) * 1988-12-08 1990-09-11 The Boeing Company Flight path angle command flight control system for landing flare

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4304375A (en) * 1979-05-17 1981-12-08 Textron Inc. Electrically controlled elevator
FR2604001B1 (en) * 1986-09-15 1988-12-09 Aerospatiale ELECTRIC FLIGHT CONTROL SYSTEM WITH INCIDENT PROTECTION FOR AIRCRAFT

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3167276A (en) * 1961-09-29 1965-01-26 Honeywell Inc Control apparatus
US3142457A (en) * 1962-07-16 1964-07-28 Boeing Co Stall pattern lift regulator for airplanes
US3942746A (en) * 1971-12-27 1976-03-09 General Dynamics Corporation Aircraft having improved performance with beaver-tail afterbody configuration
US4003533A (en) * 1973-10-01 1977-01-18 General Dynamics Corporation Combination airbrake and pitch control device
US4142699A (en) * 1977-02-04 1979-03-06 Boeing Commercial Airplane Company Lateral control system
US4261537A (en) * 1979-02-28 1981-04-14 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Velocity vector control system augmented with direct lift control
US4442490A (en) * 1980-09-26 1984-04-10 S-Tec Corporation Aircraft pitch stabilization apparatus
US4485446A (en) * 1981-09-08 1984-11-27 The Boeing Company Aircraft lift control system with acceleration and attitude limiting
US4569494A (en) * 1982-12-23 1986-02-11 The Boeing Company Pitch control of swept wing aircraft
US4744532A (en) * 1984-01-09 1988-05-17 Societe Nationale Industrielle Et Aerospatiale Aircraft flight control system
US4956780A (en) * 1988-12-08 1990-09-11 The Boeing Company Flight path angle command flight control system for landing flare

Cited By (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5375793A (en) * 1992-08-14 1994-12-27 Aerospatiale Societe Nationale Industrielle Process for the control of the control surfaces of an aircraft for the low speed compensation of a lateral path deviation
US5779191A (en) * 1996-11-12 1998-07-14 Brislawn; Mark G. Pylon flap for increasing negative pitching moments
US9764819B2 (en) 2000-02-14 2017-09-19 Aerovironment, Inc. Active dihedral control system for a torsionally flexible wing
US9120555B2 (en) 2000-02-14 2015-09-01 Aerovironment Inc. Active dihedral control system for a torisionally flexible wing
US20100308161A1 (en) * 2000-02-14 2010-12-09 Aerovironment Inc. Aircraft control system
US7802756B2 (en) * 2000-02-14 2010-09-28 Aerovironment Inc. Aircraft control system
US20080001028A1 (en) * 2000-02-14 2008-01-03 Greg Kendall Aircraft control system
US6641086B2 (en) * 2001-08-14 2003-11-04 Northrop Grumman Corporation System and method for controlling an aircraft
US20050242234A1 (en) * 2004-04-29 2005-11-03 The Boeing Company Lifters, methods of flight control and maneuver load alleviation
US7350750B2 (en) * 2004-05-28 2008-04-01 Airbus France Control system for airbrakes of an aircraft
US20050263644A1 (en) * 2004-05-28 2005-12-01 Airbus France Control system for airbrakes of an aircraft
US20090314897A1 (en) * 2004-07-16 2009-12-24 Airbus France Procedure and device for improving the maneuverability of an aircraft during the approach to landing and flare-out phases
US7896293B2 (en) * 2004-07-16 2011-03-01 Airbus Procedure and device for improving the maneuverability of an aircraft during the approach to landing and flare-out phases
US20070057114A1 (en) * 2004-07-16 2007-03-15 Airbus France Procedure and device for improving the maneuverability of an aircraft during the approach to landing and flare-out phases
US7614587B2 (en) * 2004-07-16 2009-11-10 Airbus Procedure and device for improving the maneuverability of an aircraft during the approach to landing and flare-out phases
WO2006082305A2 (en) * 2005-02-04 2006-08-10 Airbus France Method and device for piloting a pitching aircraft
US7720578B2 (en) * 2005-02-04 2010-05-18 Airbus France Method and device for piloting a pitching aircraft
US20090125165A1 (en) * 2005-02-04 2009-05-14 Airbus France Method and device for piloting a pitching aircraft
FR2881849A1 (en) * 2005-02-04 2006-08-11 Airbus France Sas METHOD AND DEVICE FOR CONTROLLING A PLANE IN BLANK
WO2006082305A3 (en) * 2005-02-04 2008-01-10 Airbus France Method and device for piloting a pitching aircraft
US8561948B2 (en) * 2005-11-29 2013-10-22 Airbus Operations Sas Method for ensuring the safety of an aircraft flying horizontally at low speed
US20080272242A1 (en) * 2005-11-29 2008-11-06 Airbus France Method for ensuring the safety of an aircraft flying horizontally at low speed
US9405295B2 (en) 2006-12-11 2016-08-02 Embraer S.A. Flight control system
US20100217460A1 (en) * 2006-12-11 2010-08-26 Polati De Souza Alvaro Vitor Flight control system
WO2008070944A3 (en) * 2006-12-11 2009-04-02 Embraer Aeronautica Sa Flight control system
US20100042271A1 (en) * 2007-03-15 2010-02-18 Torsten Holzhausen Method and device for moveable tail trimming in an aircraft
US8489257B2 (en) * 2007-03-15 2013-07-16 Airbus Operations Gmbh Method and device for moveable tail trimming in an aircraft
RU2445671C2 (en) * 2010-02-25 2012-03-20 Московский государственный университет приборостроения и информатики System for adaptive control of aeroplane on pitch angle
EP2506107B1 (en) * 2011-03-31 2020-03-11 Honeywell International Inc. System for controlling the speed of an aircraft
US20140236399A1 (en) * 2013-02-19 2014-08-21 Airbus Operations (Sas) Method and device for estimating an unwanted pitch moment of an aircraft, and applications to the pitch control of the aircraft
US9120556B2 (en) * 2013-02-19 2015-09-01 Airbus Operations (Sas) Method and device for estimating an unwanted pitch moment of an aircraft, and applications to the pitch control of the aircraft
US20160200419A1 (en) * 2014-11-12 2016-07-14 The Boeing Company Methods and apparatus to control aircraft horizontal stabilizers
JP2016153296A (en) * 2014-11-12 2016-08-25 ザ・ボーイング・カンパニーＴｈｅ Ｂｏｅｉｎｇ Ｃｏｍｐａｎｙ Methods and apparatus to control aircraft horizontal stabilizers
US9731813B2 (en) * 2014-11-12 2017-08-15 The Boeing Company Methods and apparatus to control aircraft horizontal stabilizers
CN105584627A (en) * 2014-11-12 2016-05-18 波音公司 Method and apparatus to control aircraft horizontal stabilizer
AU2015213424B2 (en) * 2014-11-12 2018-11-15 The Boeing Company Methods and apparatus to control aircraft horizontal stabilizers
EP3020630A1 (en) * 2014-11-12 2016-05-18 The Boeing Company Methods and apparatus to control aircraft horizontal stabilizers
CN105584627B (en) * 2014-11-12 2020-06-02 波音公司 Method and device for controlling a horizontal stabilizer of an aircraft
US20180101181A1 (en) * 2016-10-11 2018-04-12 Airbus Operations Sas Aircraft flight control method and system
US10386859B2 (en) * 2016-10-11 2019-08-20 Airbus Operations Sas Aircraft flight control method and system
CN114056551A (en) * 2022-01-12 2022-02-18 中国空气动力研究与发展中心低速空气动力研究所 Virtual wing belly flap and wing body fusion airplane, constant air blowing method and variable-angle air blowing method
CN114056551B (en) * 2022-01-12 2022-04-01 中国空气动力研究与发展中心低速空气动力研究所 Virtual wing belly flap and wing body fusion airplane, constant air blowing method and variable-angle air blowing method

Also Published As

Publication number Publication date
EP0454549B1 (en) 1994-02-09
DE69101154T2 (en) 1994-07-21
FR2661149B1 (en) 1992-08-14
ES2051089T3 (en) 1994-06-01
CA2040874A1 (en) 1991-10-25
CA2040874C (en) 2001-06-12
DE69101154D1 (en) 1994-03-24
FR2661149A1 (en) 1991-10-25
EP0454549A1 (en) 1991-10-30

Similar Documents

Publication Publication Date Title
US5112009A (en) Pitch control system for aircraft
US4825375A (en) Apparatus and methods for apportioning commands between aircraft flight control surfaces
US5979835A (en) Aircraft pitch-axis stability and command augmentation system
US5072893A (en) Aircraft modal suppression system
US4168045A (en) Speed and collective pitch bias of helicopter longitudinal cyclic pitch
US5330131A (en) Engines-only flight control system
US4094479A (en) Side slip angle command SCAS for aircraft
US3704843A (en) Aircraft control system
US3279725A (en) Flight controller for flexible vehicles
EP0253614A2 (en) Vertical flight path and airspeed control system for aircraft
US4797674A (en) Flight guidance system for aircraft in windshear
US5527002A (en) Electrical flight control system for an airplane, with attitude protection on takeoff
US5365446A (en) System for integrated pitch and thrust control of any aircraft
US3236478A (en) Flight control system
US3448948A (en) Aircraft speed controller
GB2136746A (en) Mechanism for increasing the flutter speeds of transonic aircraft wings
US4767085A (en) Synthetic speed stability flight control system
US4485446A (en) Aircraft lift control system with acceleration and attitude limiting
US3533579A (en) Aircraft speed controller
US4481586A (en) Adaptive washout circuit for use in a stability augmentation system
Turner Some flight characteristics of a deflected slipstream V/STOL Aircraft
US3119583A (en) Linkage in aircraft power control systems
Yamamoto Impact of aircraft structural dynamics on integrated control design
Russell et al. A Flight Investigation of the Handling Characteristics of a Fighter Airplane Controlled Through a Rate Type of Automatic Control System
Wolowicz et al. Summary of stability and control characteristics of the XB-70 airplane

Legal Events

Date Code Title Description
AS Assignment

Owner name: AEROSPATIALE SOCIETE NATIONALE INDUSTRIELLE, FRANC

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST.;ASSIGNOR:FARINEAU, JACQUES;REEL/FRAME:005951/0169

Effective date: 19911104

STCF Information on status: patent grant

Free format text: PATENTED CASE

FPAY Fee payment

Year of fee payment: 4

FEPP Fee payment procedure

Free format text: PAYOR NUMBER ASSIGNED (ORIGINAL EVENT CODE: ASPN); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

FPAY Fee payment

Year of fee payment: 8

FPAY Fee payment

Year of fee payment: 12

AS Assignment

Owner name: AIRBUS FRANCE, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:EUROPEAN AERONAUTIC DEFENCE AND SPACE COMPANY EADS FRANCE;REEL/FRAME:017575/0111

Effective date: 20051221

Owner name: EUROPEAN AERONAUTIC DEFENCE AND SPACE COMPANY EADS

Free format text: CHANGE OF NAME;ASSIGNOR:AEROSPATIALE SOCIETE NATIONALE INDUSTRIELLE;REEL/FRAME:017575/0062

Effective date: 20051221